A continuing interest exists in industry for improved gas turbine engines. For many reasons, gas turbine engines continue to be useful in a wide variety of applications. Operational costs could be substantially improved in many applications by adoption of an improved gas turbine engine that would increase operating efficiency as compared to currently utilized gas turbine designs. Further, from the point of view of maintenance costs, it would be desirable to develop improved gas turbine engine designs that reduce the mass of rotating components, since such rotating components are comparatively costly when replacement or repair becomes necessary, as compared to non-rotating parts, which although subject to stress and strain from temperature and pressure, are not subject to additional loads due to rotary motion. Thus, it can be appreciated that it would be advantageous to provide a new, high-efficiency gas turbine design which minimizes moving parts. Generally, it would be advantageous to provide more horsepower with less weight. And more specifically, it would be advantageous to provide small gas turbines having a relatively high overall efficiency, particularly in the one-half to three megawatt range.
Components of gas turbine engines include a compressor, a combustor, a nozzle or choke, and a turbine. Although supersonic compressors have been contemplated for use in gas turbine engines, in so far as we are aware, there remain various technical problems in the field and with respect to which better solutions are required in order to improve operational capability and compression efficiency. In particular, although relatively low pressure ratio supersonic compressors have been proposed for gas turbine engines, there still remains a need for gas turbine engines that can be easily started, yet be operated at high compression ratios. Further, it would be advantageous to avoid configurations that present moving shocks, such as between moving blades, or between moving and fixed blades, or between moving blades and fixed walls, in order to more simply achieve stability for normal shock location in the compressor of a gas turbine engine. Further, new designs that might reduce or minimize mechanical issues such as axial thrust loading and the requirement for expensive bearings and related equipment, would be desirable.
Although a variety of gas turbine engines with supersonic compressors have been contemplated, the work of G. F. Hausmann as reflected in U.S. Pat. No. 2,947,139, issued Apr. 2, 1960, and entitled By-Pass Turbojet is instructive of such work generally, and thus is suggestive of technical problems that remain in the field and with respect to which better solutions are required in order to improve operational capability and compression efficiency. The Hausmann patent describes the use of counter-rotating impulse blade rotors (or a single rotor in another embodiment) with a downstream stator, combustor, and convergent/divergent exhaust nozzle. However, that device is directed at aircraft propulsion applications. Flow is maintained substantially axially, with attendant losses. And, like other prior supersonic compressor designs, the design does not address minimization of the total “wetted area”—including the leading edges subject to bow wave shock losses which result in loss of efficiency. And, although the Hausmann patent suggests “high pressure” ratios, the pressure ratios noted therein are quoted as being “6 or 7 with adiabatic efficiencies which approach multi-stage subsonic compressors.”
In short, there remains a need to provide a design for a high pressure ratio—such as in excess of seven to one (7:1), or even twice that, or more—in a supersonic compressor that simultaneously resolves various practical problems seen in previous designs. Such problems include (a) providing for starting of a compressor designed for high pressure ratio operation so as to control a normal shock wave at an effective location in a supersonic diffuser designed for high pressure ratio and efficient compression, (b) avoiding excessive numbers of leading edge structures (such as may be encountered in prior art multi-bladed stators), and minimizing other losses encountered by a high velocity supersonic gas stream upon entering a diffuser, and (c) providing for effective boundary layer control, especially as related to retention of a normal shock at a desirable location, in order to achieve high compression ratios in an efficient manner. Further, improvements in overall power to weight ratio (that is, providing power in a small, compact package), and in overall cycle efficiency (thus improving fuel economy), would be a distinct advantage in many, if not most applications.